The invention relates to a device and process for controlling the attitude of a spacecraft to be rotated about an axis of its body, i.e., the axis of rotation, and more particularly to an attitude control device comprising actuators for generating torques about the axis of rotation as well as about two lateral axes which are orthogonal with respect to the axis of rotation and to one another. Sensors are provided for the formation of angular velocity signals with respect to the three axes. Two modulators, which are in each case connected in front of the actuators assigned to one of the two lateral axes emit controlling signals to the actuators and have a variable dead zone. Two regulator networks, which receive one of the two lateral-axis angular velocity signals respectively, furnish a control signal for one of the two modulators and have a first signal path as well as a second signal path which is connected in parallel to the first one and contains an integrator.
This type of a device and this type of a process are known from U.S. Pat. No. 4,725,024. The object of U.S. Pat. No. 4,725,024 is to send a three-axis stabilized satellite, which is on a low, almost circular orbit, into an elliptic transfer path by igniting a perigee power unit. The apogee of this transfer path coincides with the radius of the endeavored geostationary orbit. Before igniting the perigee engine, for reasons of stability, the satellite must be rotated about a body axis of rotation which is to coincide with the thrust vector of the thrust exercised by the perigee power unit. In this case, the perigee power unit is still docked to the satellite, and the axis of symmetry of the power unit is congruent with an axis of symmetry of the satellite. In the case of U.S. Pat. No. 4,725,024, this is the roll axis. As known, the roll axis is one of the body's three axes of the satellite forming a rectangular system of coordinates, which also include the yaw axis as well as the pitch axis as the lateral axes. In the final operating condition of the satellite on the geostationary orbit, the roll axis must be oriented in the flight direction; the yaw axis must be oriented toward the center of the earth; and the pitch axis must be oriented perpendicularly to the two as well as to the plane of the orbit.
The attitude control system of a three-axis stabilized satellite includes a number of actuators, for example, fuel nozzles, which are capable of furnishing in a targeted manner torques or controlling torques about the three above-mentioned axes. The known attitude control device according to U.S. Pat. No. 4,725,024 also has sensors in the form of gyroscopes which furnish angular velocity signals with respect to rotations about the three axes of the body. Each of the two lateral-axis angular velocity signals is fed into a regular network which furnishes a control signal for a modulator which comprises a dead zone and, in turn, has the task of generating discrete control signals for the actuators assigned to the respective lateral axis. The two regulator networks have a first signal path as well as a second signal path which is connected in parallel to the first and contains an integrator. The two signal paths are fed together in front of the assigned modulator in a summation element.
Before being discharged from the space tug, the satellite with the docked perigee power unit already rotates slowly at approximately two revolutions per minute about the axis of rotation. After the discharge of the thus formed spacecraft from the loading bay of the space tug, via the actuating of the corresponding actuators, the rotation about the axis is continuously increased to a value of approximately forty revolutions per minute. If possible, care must be taken during this spin-up phase that the inevitably occurring nutation is controlled to a constant amplitude and the direction of the axis of rotation in the inertial space is maintained as precisely as possible.
The excitation of nutation vibrations may have different causes. Thus, generally, because the mass distribution is not completely symmetrical, the main inertia axis of the satellite/perigee power unit combination will not coincide with the original axis of rotation. The rotation is to be stabilized about this main inertia axis which represents the axis with the smallest moment of inertia. However, the actuators causing the rotation are arranged symmetrically with respect to the geometric axis of rotation so that lateral-axis torques already occur here. Even higher lateral-axis torques are caused by the fact that the center of gravity of the satellite/perigee power unit combination shifts considerably with respect to the center of gravity of the satellite. The actuators serving the attitude control, in their arrangement on the satellite as well as with respect to the direction of action of their controlling torques, are adapted to a constellation in the case of which the satellite has already separated from the perigee power unit. Since however, the rotation caused during the spin-up phase as well as the then required attitude control are to be carried out exclusively via the actuators which anyhow exist in the normal operation, the shifting of the center of gravity when an actuator is operated causes considerable environmental torques about other axes. In addition, an increase in amplitude of the nutation may be caused by a sloshing of liquid fuel.
In the case of the attitude control device according to U.S. Pat. No. 4,725,024, it is provided that, after approximately half the spin-up operation, the dead zones of the modulators must suddenly be expanded. This has the purpose of permitting larger nutation amplitudes without any intervention of the power units. However, since the destabilizing influences increase with the nutation amplitude, a control of the nutation to a constant amplitude would be extremely expedient. However, by means of the known attitude control device, this object cannot be achieved, particularly since the regulator networks operate in the two signal paths with a respective constant amplification.
It is a further disadvantage of the known attitude control device that the gyroscope used for measuring the angular velocity about the axis of rotation will be saturated after a short period of time because it is designed only for relatively low angular velocities occurring during normal operation. As a result, the angular position information for all three axes is lost shortly after the start of the spin-up phase, as well as the angular velocity information for the axis of rotation. There is the danger of an undesirable drifting-away of the axis of rotation from the inertial orientation that is to be maintained, if possible. Finally, the lateral-axis angular velocity signals contain constant signal fractions which result in an asymmetrical controlling-out of the dead zones of the modulators and, thus, to the triggering of one-sided control interventions as well as to the adjustment of the axis of rotation. The constant signal fractions are caused by the fact that, as mentioned above, the spacecraft rotates about the main axis of inertia which has unavoidable deviations from the geometric axis. The gyroscopes, serving for measuring the angular velocities, are aligned to the geometric satellite axes so that the angular velocity signals contain fractions which are proportional to the variation of the compass between the geometric and the main inertia axis and the angular velocity about the axis of rotation.
The invention is therefore based on the object of providing a device of the above-mentioned type which mainly permits the limiting of the nutation amplitude to a constanct value in a reliable manner. According to the invention, this object is achieved by an attitude control device comprising actuators for generating torques about the axis of rotation as well as about two lateral axes which are orthogonal with respect to the axis of rotation and to one another. Sensors are provided for the formation of angular velocity signals with respect to the three axes. Two modulators, which are in each case connected in front of the actuators assigned to one of the two lateral axes emit controlling signals to the actuators and have a variable dead zone. Two regulator networks, which receive one of the two lateral-axis angular velocity signals respectively, furnish a control signal for one of the two modulators and have a first signal path as well as a second signal path which is connected in parallel to the first one and contains an integrator. Two first multiplication elements are each connected into one of the two first signal paths and use factors which are dimensioned proportionally to the rotation axis angular velocity signal. Two second multiplication elements are each connected into one of the two second signal paths behind the integrator and use factors which are dimensioned proportionally to the square of the rotation axis angular velocity signal. Devices are provided for varying the thresholds of the dead zone of the respective modulator proportionally to the square of the rotation axis angular velocity signal.
Other objects, advantages and novel features of the present invention will become apparent from the following detailed description of the invention when considered in conjunction with the accompanying drawings.